Alloy composition

ABSTRACT

An alloy composition consisting essentially of, by atomic percent, between 6.5 and 8% of elements selected from the group consisting of aluminium, titanium, tantalum and niobium, between 37 and 47% of elements selected from the group consisting of chromium, cobalt, iron, molybdenum and tungsten, and between 1.8 and 3.8% of elements selected from the group consisting of tungsten, molybdenum, niobium and tantalum, and, optionally, by weight percent, up to 0.04% carbon, up to 0.07% boron, and up to 0.1% zirconium, the balance being nickel and incidental impurities.

FIELD OF THE INVENTION

The present invention relates to a nickel based alloy composition and agas turbine engine component comprising a nickel based alloy.

BACKGROUND TO THE INVENTION

FIG. 1 shows a high by pass ratio gas turbine engine 10. The engine 10comprises an air intake 11 and a propulsive fan 12 that generates twoairflows A and B. The gas turbine engine 10 comprises, in axial flow A,an intermediate pressure compressor 14, a high pressure compressor 16, acombustor 18, a high pressure turbine 20, an intermediate pressureturbine 22, a low pressure turbine 24 and an exhaust nozzle 26. Anacelle 30 surrounds the gas turbine engine 10 and defines, in axialflow B, a bypass duct 13.

The combustor 20 is shown in further detail in FIG. 2. The combustorcomprises a combustor casing 32, within which is located a metalcombustor liner 34. The combustor liner 34 is in turn covered incombustor liner tiles (not shown), which are made of a ceramic material.In use, air and fuel flow into the combustor 20, where the fuel isburned, producing hot combustion gases.

The combustor liner 34 must operate at high temperature in excess of800° C. (and perhaps as high as 900° C.) for long periods of time.Higher combustion chamber temperature will result in higher thermalefficiencies of the gas turbine engine, and so this temperature must bemade as high as possible through the use of high temperature alloys.High temperature alloys are also used in other parts of the engine, suchas in the turbines and exhaust duct.

In order to permit operation at high temperatures, and provide a longservice life, suitable alloys must also have a number of otherproperties, in addition to a high temperature capability. For example,they must have a high ultimate tensile strength, yield strength, stressrupture resistance, ductility, stability at high temperatures,resistance to thermal stresses, density and environmental resistance(e.g. resistance to hot corrosion and oxidation). In the art, the“stability” of an alloy is normally understood to refer to the alloy'spropensity to precipitate detrimental phases (i.e. an alloy having ahigh stability will have a low propensity to precipitate detrimentalphases). An example of a detrimental phase is the sigma (a) phase, whichcan occur when the alloy is subjected to high temperatures for extendedtime periods (known as “dwell”).

Table 1 below defines prior nickel based compositions suitable for usein combustor liners for gas turbine engines. All amounts are given inweight percentages.

Alloy described in Alloy described in U.S. Pat. No. U.S. Pat. No.Composition 4,174,213 4,080,201 Nickel (Ni)  42-70% Balance Chromium(Cr)  15-35% 12-18%  Cobalt (Co) 0.1/10%  <2% Iron (Fe) 7.5-35%  0-3%Manganese (Mn)    <2% — Tungsten (W) 0.1-10%  0-7% Niobium (Nb) 0.05-1%— Tantalum (Ta) — <0.75%  Silicon (Si)    <2% 0.08% Aluminium (Al) — 0.5% Titanium (Ti) 0.05-1% 0.75% Carbon (C) 0.03-0.2%  0.02% Molybdenum(Mo) 4.5-15% 10-18% 

Other alloys used in combustor liners include Haynes 188™, Haynes 230™and Nimonic 263™.

It is also desirable for the alloy composition to have a low cost (interms of the elemental cost of the alloy), and low density (particularlywhere the combustor is for use in aerospace gas turbine engines) and besuitable for low cost production methods such as casting, wroughtprocessing, powder metallurgy or direct laser deposition. It is alsodesirable that the final alloy is highly weldable (i.e. does notmicro-segregate when melted), particularly where the alloy is to be usedin direct laser deposition. Alloys having the above properties, andtherefore being suitable for use in gas turbine engine components suchas combustor liners, are generally known in the art as “superalloys”,and are sometimes also referred to as “high performance alloys”.

The present invention seeks to provide an improved alloy composition andan improved gas turbine engine component which solves some or all of theabove problems.

SUMMARY OF THE INVENTION

According to a first aspect of the present invention there is providedan alloy composition consisting essentially of, by atomic percent,between 6.5 and 8% of elements selected from the group consisting ofaluminium, titanium, tantalum and niobium, between 37 and 47% ofelements selected from the group consisting of chromium, cobalt, iron,molybdenum and tungsten, and between 1.8 and 3.8% of elements selectedfrom the group consisting of tungsten, molybdenum, niobium and tantalum,and, optionally, by weight percent, up to 0.04% carbon, up to 0.07%boron and up to 0.07% zirconium, the balance being nickel and incidentalimpurities.

It has been found that the alloy composition of the present inventionhas a high strength at temperatures at around 900° C., and is alsorelatively lightweight, resistant to environmental degradation,inexpensive and suitable for conventional forming and machiningprocesses. These properties make the alloy particularly suitable for usein gas turbine engine components such as combustor linings.

According to a second aspect of the invention, there is provided acomponent of a gas turbine engine formed of an alloy in accordance withthe first aspect of the invention.

According to a third aspect of the invention there is provided a gasturbine engine comprising a component according to the second aspect ofthe invention.

Accordingly, the invention provides a gas turbine engine having one ormore components capable of operation at high temperatures. Consequently,the gas turbine engine can operate at a higher thermal efficiency, andhence lower specific fuel consumption, or may have a longer service lifebetween overhauls, resulting in a lower operating cost.

Further features of the invention are described in the attached claims.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the invention are described and shown in the followingdrawings, in which:

FIG. 1 shows a cross sectional view of a gas turbine engine;

FIG. 2 shows a cross sectional view of part of the engine of FIG. 1;

FIG. 3 is a graph showing various properties of an alloy in accordancewith the present disclosure plotted against temperature, compared to aprior alloy;

FIG. 4 is a graph showing the maximum stress of both an alloy inaccordance with the present disclosure and prior alloys, plotted againsttemperature;

FIG. 5 is a graph showing the proportion of various phases in an alloyin accordance with the present disclosure, plotted against temperature;

FIG. 6 is a graph showing the chromium activity of an alloy inaccordance with the present disclosure compared with that of Nimonic263™, plotted against temperature; and

FIG. 7 is a scanning electron microscope image of a sample of an alloyin accordance with the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 shows a high by pass ratio gas turbine engine 10. The engine 10comprises an air intake 11 and a propulsive fan 12 that generates twoairflows A and B. The gas turbine engine 10 comprises, in axial flow A,an intermediate pressure compressor 14, a high pressure compressor 16, acombustor 18, a high pressure turbine 20, an intermediate pressureturbine 22, a low pressure turbine 24 and an exhaust nozzle 26. Anacelle 30 surrounds the gas turbine engine 10 and defines, in axialflow B, a bypass duct 13.

The combustor 20 is shown in further detail in FIG. 2. The combustorcomprises a combustor casing 32, within which is located a metalcombustor liner 34. The combustor liner 34 is in turn covered incombustor liner tiles (not shown), which are made of a ceramic material.In use, air and fuel flow into the combustor 20, where the fuel isburned, producing hot combustion gases.

Table 2 shows a compositional range of an alloy in accordance with thedisclosure, which is suitable for one or more components of the gasturbine engine 10 (and particularly suitable for use as the material ofthe combustor liner 34):

TABLE 2 wt. % target maximum minimum Ni Balance Balance Balance Cr20.0-20.4 22 18.0 Co  9.3-11.3 12 8.5 Mo  0-0.2 0 0.5 Fe 7.4-8.4 9.4 6.4Mn 0.2-0.4 1.5 0.0 W 4.2-4.6 5.3 3.5 Nb 1.0-1.2 1.5 0.0 Ta 0.6-0.8 1.00.4 Si 0.2-0.4 1.5 0 Al 1.8-2.0 2.3 1.5 Ti 1.6-1.8 2.4 0.5 C 0.015-0.0250.04 0.005 B 0.015-0.025 0.04 0.005 Zr 0.05-0.07 0.10 0.02

Oxygen may also be present, in the form of surface oxides. Otherincidental impurities may also be present in the alloy. In general,other impurities should be kept to a minimum, in particular sulphurousbased impurities.

Various compositions can be produced using the maximum and minimumelemental quantities described in table 2. It has been found that allcompositions within the bounds of the maximum and minimum quantities oftable 2 result in alloys which have acceptable properties for use as acombustor liner 34.

The narrower target compositional range shown in table 2 has improvedqualities over the alloy compositions lying outside this range, butwithin the wider range shown in the maximum and minimum columns. Theseminimum and maximum amounts are based upon sensitivity studies using acomputational materials prediction tool which show the deviation ofpredicted properties with changing composition.

Table 3 below shows a nominal composition in accordance with the presentdisclosure (composition 1). The physical properties of this compositionare described in further detail below.

TABLE 3 wt. % Composition Ni Bal. Cr 20.2 Co 10.3 Mo 0.1 Fe 7.9 Mn 0.3 W4.4 Nb 1.1 Ta 0.7 Si 0.3 Al 1.9 Ti 1.7 C 0.015-0.025 B 0.015-0.025 Zr0.06

Again, oxygen may also be present, in the form of surface oxides.Similarly, impurities may also be present.

The composition comprises nickel, which forms a continuous matrixcomprising a face centred cubic (FCC) nickel based austenitic phasecontaining solid solution elements. The nickel based continuous matrixis known within the art as the “gamma (γ)” phase. Some of the alloyingelements form a primary strengthening phase known as a “gamma prime(γ′)” phase in an amount such that the volume fraction of the γ′ isapproximately 20% at 900° C. The γ′ phase has the general formula Ni₃x,where x comprises elements selected from titanium, aluminium, tantalumand niobium, and usually comprises an ordered intermetallic L₁₂ crystalstructure. Formation of the γ′ phase occurs in the solid state as thesupersaturated solid solution of γ-nickel is cooled below its solvustemperature. Other elements (such as cobalt, iron, and tungsten) providesolid solution strengthening within the nickel matrix.

Some of the elements given in tables 2 and 3 partition to the gammaprime phase within the alloy. In this case, each of aluminium, titaniumtantalum and niobium partition to the gamma phase. Though the relativeamounts of each of these elements may vary between compositions withinthe scope of the present disclosure, the total amount of elements thatpartition to the gamma prime phase is between 6.5 and 8.5 atomicpercent. This has been found to contribute to the desirable propertiesof the disclosed alloy, in particular, ultimate tensile strength.

Some of the elements given in tables 2 and 3 partition to the gammaphase. Each of chromium, cobalt, iron, molybdenum and tungsten partitionto the gamma phase. Again, the relative amounts of each of theseelements may vary between compositions within the scope of the presentdisclosure. However, the total amount of elements that partition to thegamma phase is between 37 and 47 atomic percent.

The alloy includes the following refractory (i.e. high meltingtemperature) alloys, which offer significant strengthening in the alloyat the temperatures at which gas turbine engines operate: tungsten,molybdenum, niobium and tantalum. These refractory alloys are includedin amounts between 1.8 and 3.8 atomic percent.

The chromium present in the composition is required to maintain aprotective oxide scale on the surface of the billet (or finishedcomponent), providing resistance to oxidation, type I and II type hotcorrosion, and dwell fatigue crack resistance. If the chromium contentis too high (i.e. significantly above (20%), then the formation ofdeleterious phases is encouraged, which will impair the mechanicalproperties of the alloy.

Cobalt is added to provide additional solid solution strengthening tothe gamma matrix and reduce the stacking fault energy. Cobalt can beadded in comparatively large quantities (up to 12%) due to itscompatibility with nickel. However, too much cobalt (i.e. significantlymore than 12%) will increase the propensity to form deleterious phasesin the alloy at the temperatures at which gas turbine engine componentstypically operate.

Iron has good solubility within the gamma phase, and is added as a solidsolution strengthener. It has the additional benefits that it is lowcost and decreases the density of the alloy, resulting in a lightweight,low cost alloy. However, an iron content that is too high (i.e.significantly above 10%) will promote the formation of the undesirableLaves phase at the temperatures at which gas turbine engines operate.

Tungsten is also added for its solid solution strengthening properties.Tungsten is thought to be a more potent solid solution strengthener thaneither cobalt or iron, but cannot be added in large quantities (i.e. atamounts significantly above 5.5%) due to the increased promotion ofdeleterious intermetallics, and its adverse effect upon the alloy'sdensity. The amount of tungsten present in the composition is unusuallyhigh for a high temperature nickel based alloy, and has been found togreatly contribute to the high ultimate tensile strength of the alloy athigh temperatures.

Niobium will partition to the gamma prime phase and providestrengthening, resisting the movement of dislocations through the gammaprime phase. However, the addition of too much niobium (i.e.significantly more than 1.5%) will result in the precipitation of thedeleterious delta phase at the temperatures at which gas turbinecomponents operate, which is not desired in the present invention.Niobium is also a potent metal carbide former (in conjunction with thecarbon present in the alloy), which will improve dwell crack properties.

Tantalum is an effective gamma prime strengthener, preventing themovement of dislocations, which will give improved creep and othermechanical properties. The addition of tantalum will significantlyincrease the density and elemental cost of the alloy, which are bothundesirable. However, unlike the other refractory metals, tantalum canbe added in greater quantities before the alloy stability is compromised(up to 1%). Tantalum is also a potent metal carbide former.

Silicon promotes the formation of a more stable and resistive oxidescale than chromium alone. However, its introduction leads to greatinstability in anything other than small quantities (i.e more than1.5%), which has a significant impact upon the mechanical properties ofarticles formed from the alloy. While a large quantity of siliconsuppresses formation of the deleterious eta phase, too much promotes theformation of the deleterious G phase.

Molybdenum is optionally present in the alloy. Molybdenum is postulatedto have a positive effect on the environmental resistance of the alloy.It will also act as a solid solution strengthener but is more prone topromoting the formation of deleterious phases then other solid solutionstrengthening elements.

Aluminium is essential for promoting the formation of the gamma primephase, which provides the major strengthening mechanism for the alloy.The control of its quantity is crucial to achieve the correct balance ofproperties (particularly in terms of the ratio of aluminium. Too muchaluminium, and the alloy will be unprocessible (i.e., difficult toweld)—too little, and the alloy will have insufficient mechanicalstrength. In addition, aluminium will improve the oxidation resistanceand lower the density of the alloy, which are both highly importantconsiderations in aerospace gas turbine engine components.

Titanium will strengthen the gamma prime phase as well as increase thefraction of gamma prime present. It will also reduce the density of thealloy. The addition of too much titanium (i.e. significantly more than2.4%) will promote the formation of the deleterious eta phase, promotethe formation of too much gamma prime, and may compromise theenvironmental resistance of the alloy by increasing oxide thickeningrates.

The aluminium to titanium ratio is generally greater than 1:1. This,along with a Cr/Ti ratio greater than 10:1, promotes a good oxidationand corrosion resistance. However, it is still recognised thesignificant strengthening benefit that Ti adds and unlike most otherstrengthening elements, it has no density penalty.

Carbon, boron and zirconium are added in small amounts (as shown intable 2) to form carbides and borides on the grain boundaries, whichstrengthen the grain boundaries of the alloys. Their amounts have beenempirically optimised to prevent crack dwell fatigue and also preventmelt anomalies which will improve weldability.

The described alloy compositions can be used in various componentmanufacture methods, such as any of powder metallurgy methods, castingor laser deposition welding. The described alloy compositions areparticularly suitable for laser deposition welding, since the alloy ishighly stable, and does not tend to microsegregate when melted.

To generate a forging having the required balance of properties, it maybe necessary to subject the forging to a heat treatment process. Thisheat treatment may be performed either above or below the gamma primesolvus temperature to obtain the desired gamma prime precipitatedistribution. Optionally, this may be followed by an ageing treatment,which nominally would be 4-16 hours at 850-900° C.

Various material properties of composition 1 are shown in FIGS. 3 to 6.

FIG. 3 shows, plotted against temperature, the rupture stress, ultimatetensile stress, yield stress and design space (i.e. the minimum of therupture stress, ultimate tensile stress (UTS), and yield stress) ofcomposition 1 (labelled as “optimised combustor” on the graph), comparedto equivalent values of Nimonic 263™. The design space essentiallyrepresents the maximum stress that can be applied to the alloy prior toa failure of some sort at a given temperature.

As can be seen, composition 1 has a higher predicted yield stress andstress rupture behaviour than Nimonic 263™, but a lower predicted UTS.

FIG. 4 compares the design space of composition 1 (again labelled as“optimised combustor alloy”), compared to Nimonic 263™, Haynes 282,Haynes 230 and Haynes 188. As can be seen, the design space of alloycomposition 1 of the present disclosure is higher than all of the prioralloys at lower temperatures (less than 600° C.), and only lower thanHaynes 282 at higher temperatures.

However, it is thought that at these higher temperatures, the thermalstresses in Haynes 282 will be higher than those in composition 1.Consequently, the alloy of the present disclosure will be able tooperate at higher temperatures than even Haynes 282.

The merit of an alloy, P to resist thermal stresses can be given as thefollowing merit index:

$P = \frac{\sigma_{YS}}{E\; \rho \; \alpha}$

Where σ_(YS) is the yield stress, E is the Young's modulus, ρ is theresistivity and α is the thermal expansivity. Accordingly, the alloy ofthe present invention is predicted to have a high resistance to thermalstress.

FIG. 5 shows the relative molar fractions of various phases incomposition 1, as plotted against temperature. As can be seen, attemperatures around 900° C., substantially only gamma and gamma primeare present. No precipitation of sigma (σ) is observed, demonstratingthat the desired level of microstructural stability has been achievedfor use as an alloy that can be formed by direct laser deposition, andused as a combustor liner material.

FIG. 6 shows the activity of chromium of the alloy of composition 1,plotted as a function of temperature, in comparison to that of Nimonic263. As can been seen, the chromium activity of the alloy of composition1 is significantly higher, resulting in better environmental resistance(i.e. approximately half the oxidation rate of Nimonic 263).

FIG. 7 shows an image from a scanning electron microscope of the alloyof composition 1 after is has been subject to temperatures of 750° C.for 1000 hours. As can be seen, there is substantially nomicrostructural segregation or cracking. Consequently, the alloy isresistant to high dwell temperatures.

Primarily this alloy is intended to be manufactured using direct laserdeposition, however it is also suitable for other commonly usedmanufacturing techniques such as casting, powder processing and welding.

Consequently, the disclosed compositions describe alloys which possess anumber of advantageous properties compared to prior alloys. In summary,the main advantages of the new alloy in comparison to the prior art are:

1. better mechanical properties, particularly at temperatures above 800°C.;2. higher stablility with respect to the formation of deleteriousphases;3. lower elemental cost and lower density;4. lower thermal stress;5. higher chromium activity leading to superior corrosion and oxidationproperties;6. properties that better suit all of the requirements of a combustorliner alloy; and7. higher amenability to thermomechanical processing and additivemanufacture.

Although the description refers to the described alloys as beingparticularly useful for forming combustor linings for gas turbineengines, the alloys could also be used to form other components.

Although the alloy is described as being subjected to heat treatment,other material processing methods could be used to generate articlesfrom the alloy compositions having the required properties.

1. An alloy consisting essentially of, by weight percent, between 18 and20% chromium, between 8.5% and 12% cobalt, between 6.4 and 9.4% iron, upto 1.5% manganese, between 3.5 and 5.3% tungsten, up to 1.5% niobium,between 0.4% and 1% tantalum, up to 1.5% silicon, between 1.5 and 2.3%aluminium, between 0.5 and 2.4% titanium, between 0.005 and 0.04%carbon, between 0.005 and 0.07% boron, and between 0.02 and 0.10%zirconium, the balance being nickel save for incidental impurities. 2.An alloy according to claim 1, wherein the alloy consists essentiallyof, by weight percent, between 18 and 20% chromium, between 8.5% and 12%cobalt, between 6.4 and 9.4% iron, up to 1.5% manganese, between 3.5 and5.3% tungsten, up to 1.5% niobium, between 0.4% and 1% tantalum, between0.1% and 1.5% silicon, between 1.5 and 2.3% aluminium, between 1 and2.4% titanium, between 0.005 and 0.04% carbon, between 0.005 and 0.07%boron, and between 0.02 and 0.10% zirconium, the balance being nickelsave for incidental impurities.
 3. An alloy according to claim 1,wherein the alloy comprises between 0.1% and 1.5% by weight silicon. 4.An alloy composition according to claim 1, wherein the alloy consistsessentially of, by weight percent, between 20 and 20.4% chromium,between 9.3% and 11.3% cobalt, between 7.4 and 8.4% iron, between 0.2and 0.4% manganese, between 4.2 and 4.6% tungsten, between 1 and 1.2%niobium, between 0.6% and 0.8% tantalum, between 0.2 and 0.4% silicon,between 1.8 and 2.0% aluminium, between 1.6 and 1.8% titanium, between0.015 and 0.025% carbon, between 0.015 and 0.025% boron, and between0.05 and 0.07% zirconium, the balance being nickel save for incidentalimpurities.
 5. An alloy composition according to claim 4, wherein thealloy consists essentially of by weight percent, 20.2% chromium, 10.3%cobalt, 7.9% iron, 0.3% manganese, 4.4% tungsten, 1.1% niobium, 0.7%tantalum, 0.3% silicon, 1.9% aluminium, 1.7% titanium, between 0.015 and0.025% carbon, between 0.015 and 0.025% boron, and between 0.05 and0.07% zirconium, the balance being nickel save for incidentalimpurities.
 6. An alloy according to claim 1, wherein the incidentalimpurities include oxygen.
 7. An alloy according to claim 1, wherein theratio by weight of titanium to aluminium is less than 0.5.
 8. An alloyaccording to claim 1, wherein the ratio by weight of chromium totitanium is greater than
 10. 9. A gas turbine engine component formed ofan alloy in accordance with claim
 1. 10. A gas turbine engine componentaccording to claim 8, wherein the component comprises a combustor liner.11. A gas turbine engine comprising a component according to claim 8.